1. Field of the Invention
The present invention is a constant-volume, Reciprocating Thrust Valve (RTV), bipropellant rocket engine. The motor provides improved reliability and fuel efficiency compared to conventional constant pressure rocket motors for precision high-performance pulse-mode propulsion and multi-use in-space thruster applications.
2. Description of Related Art
Conventional liquid bipropellant thrusters are designed based on a constant-pressure combustion cycle. This type of thruster is not well suited for high-precision control of large spacecraft, which requires a pulse-mode operation. Precision control often requires discrete impulse bits having very short on times (<<100 ms), with relatively long off times (>1 sec). During such operation, thrusters rarely, if ever, reach steady-state operating conditions. For bipropellant thrusters, transient phenomena associated with startup and shutdown play a critical role in pulse-to-pulse performance levels and repeatability. As the pulse length becomes shorter and the duty cycle is decreased, the effects of these transients become more pronounced, resulting in drastically reduced specific impulse. Furthermore, the cumulative effects of such transients and low-yield impulse bits can have a serious and adverse impact on spacecraft system performance, reliability, and safety, including the potential for catastrophic failure due to propellant accumulation in the valves and combustion chamber.
The primary cause of pulse-mode performance degradation in constant-pressure thrusters is poor atomization and mixing of propellants resulting in highly inefficient combustion, and in some cases, no combustion at all. For very short pulses, a large percentage of propellant flows out of the combustion chamber without mixing and reacting—an unavoidable feature of constant-pressure combustion that is exacerbated by large dribble volumes, dribble volume mismatch, propellant flash vaporization, and certain combustion chamber geometrical parameters. This low-efficiency combustion process not only reduces performance, but also results in a significant potential for spacecraft contamination. In order to optimize pulse-mode performance of attitude control thrusters, designers typically seek to minimize propellant dribble volume, which requires close coupling of the valves to the injector. During steady state operation, a thermal balance is achieved between the flowing propellants, thermal radiation, and liquid-film cooling (if applicable) and the hot-combustion gases such that the valve temperature does not exceed a maximum specified temperature. During pulse-mode operation at certain duty cycles, however, this thermal balance becomes unstable due to fuel rich combustion and excessive thermal soak-back that can result in overheating of the valves.
An alternative to constant pressure combustion method is constant-volume combustion. Idealized analysis has shown that, for the same propellant supply pressure, higher performance can be obtained in a constant-volume combustion device than the traditional constant pressure rocket. In a constant-volume combustion cycle, propellants are injected into a closed-volume chamber at some low initial pressure and temperature. This closed-volume approach allows for precision-timing control of subsequent mixing, ignition, combustion and flow processes, unlike conventional constant-pressure thrusters where much of the propellant escapes the combustion chamber prior to ignition and complete combustion. When maximum pressure (complete combustion) has been achieved, a Reciprocating Thrust Valve RTV is retracted from the throat and the high-temperature combustion products escape through the nozzle to produce a single impulse bit. After the combustion chamber has been evacuated, the RTV closes and the thruster is ready for the next cycle. For a properly designed constant-volume thruster, near-ideal characteristic velocity (c star) can be achieved for a wide range of impulse bits. Furthermore, passive valving devices (e.g. reed-type valves) may be installed at the injector face to keep propellant dribble volumes full during sustained pulse-mode operation, thereby improving thruster response characteristics. This also enables the decoupling of the propellant flow control valves from the chamber to mitigate problems associated with high heat soak-back during pulse mode operation.
Unlike conventional constant-pressure hypergolic thrusters, where only the flow of propellants can be controlled, an RTV thruster allows control over parameters such as pulse repetition rate, pulse width modulation that can be optimally tuned for enhanced system operation. Given that combustion occurs in a confined volume, this design is ideally suited for scalability. Engines can be made with thrust ranges from millinewton to kilonewton.
The RTV cycle may be configured with multiple chambers firing sequentially through a single nozzle to produce quasi-continuous thrust. Alternatively, multiple thrusters can be configured in groupings or arrays to provide both main propulsion and attitude control in a single propulsion system.
Pulse Detonation motors and processes, such as those disclosed in U.S. Pat. Nos. 5,579,633; 6,062,018; 6,442,930; 6,526,936; 6,886,325; 6,931,833; 7,047,724; 2004/0050038; 2004/0154304; and 2005/0279083 and incorporated by reference, are designed to approximate or mimic constant volume combustion but are distinct from constant-volume combustion in several respects. Pulse detonation occurs in an open chamber while constant volume combustion occurs in a sealed chamber. Pulse detonation relied on extremely rapid fuel consumption (detonation) to complete combustion before unburned fuel leaves the chamber. This embodiment, on the other hand, implements a constant-volume combustion cycle in a deflagrative combustion process without the complexities of detonation. The major difficulty with a pulse detonation engine is initiating the detonation itself. The typical solution is to employ a Deflagration-to-Detonation Transition (DDT). This involves the initiation of a high-energy deflagration that accelerates down the combustion chamber until it becomes fast enough to transition to a detonation. A key difficulty in pulse detonation engines is achieving DDT without requiring an impractically long and drag-imposing combustion chamber on the vehicle. Other difficulties include, noise reduction and damping of the severe vibration caused by the operation of the engine. The present invention is not a pulse detonation motor but a deflagrating constant-volume rocket motor. The RTV mechanism provides many of the same benefits as a pulsed-detonation engine without the need for highly unpredictable detonations and complicated acoustic tuning.